The present invention relates generally to spacecraft attitude control systems and methods, and more particularly, to a closed loop spacecraft orbit control system and method.
The assignee of the present invention manufactures and deploys spacecraft that orbit the Earth. Most conventional longitude control schemes perform longitude control when the longitude error reaches a threshold value. It would be desirable to improve upon the control capability of conventional longitude control schemes.
More particularly, in the current state-of-the-art, orbit control is accomplished by monitoring orbit error and then performing maneuvers whenever the orbit errors exceed a predetermined threshold. This results in stationkeeping maneuvers that are performed weeks apart from each other. However, when developing stationkeeping strategies that use electric thrusters, it was discovered that the heritage method used by the assignee of the present invention had several shortcomings.
First, to minimize the size of solar arrays and spacecraft batteries, it was necessary to perform electric thruster stationkeeping daily. Also, since electric thrusters provide a lower thrust than conventional bipropellant thrusters, daily maneuvers were also needed to prevent long-duration maneuvers whose thrust is applied far from the orbital node which would decrease stationkeeping efficiency. Therefore, stationkeeping was performed at irregular predetermined times instead of when a threshold was tripped.
Initially, it was anticipated that daily stationkeeping could be accommodated by abandoning the heritage thresholds and performing equal amounts of orbital correction every day. While such daily stationkeeping planning performed well when the only disturbances were natural highly-predictable disturbances, it was discovered that unpredictable orbital disturbances caused daily stationkeeping planning to be performed poorly.
In particular, the electric thruster stationkeeping introduced several new disturbances to the orbit. First, to prevent the electric thruster plume from degrading the surface of the solar array, the electric thrusters were angled away from the orbit normal, introducing a radial component which affects the orbit longitude. This radial thrust produces a longitudinal shift which is compensated for by increasing the size of the orbit (above synchronous). Since the location and duration of north-south bums change over the seasons, the effects of these radial disturbances continually change and thus the necessary orbital size continually changes. Second, since the electric thrusters were gimbaled to allow momentum dumping (as described in U.S. Pat. No. 5,349,532, for example), the gimbaling changed the thruster vector direction, introducing a disturbance to the orbit. Third, the enlarged orbit resulted in a reversed longitudinal shift if the north-south burns were bypassed for a day to accommodate ranging or other satellite operations.
It is therefore an objective of the present invention to provide for an improved closed loop spacecraft orbit control system and method that improves upon conventional systems and methods.
To accomplish the above and other objectives, the present invention provides for a closed loop spacecraft orbit control system and method that provides improved orbit control of a spacecraft. An exemplary orbit control scheme implemented in the system and method comprises the following steps.
The current orbit of the spacecraft is determined. From this orbit and a control orbit, desired changes of orbit parameters are determined. Thruster firings that accomplish the desired orbit parameter changes are then determined by adding the expected orbital disturbance (if any) to a sum of terms proportional to the orbital errors. The thrusters are then fired to perform spacecraft stationkeeping and provide spacecraft orbit control.
The present invention may be used with any type of orbit determination procedure (such as using inputs derived from a global positioning system (GPS) receiver or from ranging techniques) and any type of stationkeeping algorithm that calculates thruster firings from desired orbit-parameter corrections.
The present invention thus provides for a compensation system or compensator that implements closed-loop control of the orbit parameters of a spacecraft. Since the compensator requires very few computations, it can be performed either on-board the spacecraft or in ground-based software. While conventional orbit control systems and methods perform orbit control when the orbital parameter errors reaches a threshold value, the present invention performs orbit control at regular, predetermined times.
Performing longitude control at regular times allows the longitude of the spacecraft to be controlled to tighter values. Using a linear feedback control law makes the control system and method more robust to unmodeled disturbances and improves the control stability.
The present invention preferably uses discrete linear-feedback control to determine the amount of orbit-parameter correction. The calculation of orbit-parameter correction is separated from the calculation of thruster forces or other actuation, allowing the present invention to used with any stationkeeping algorithm that calculates thruster firings from desired orbit-parameter changes.
It is worth noting that another possible method for orbit maintenance is described in U.S. Pat. Nos. 5,528,502 and 5,687,084. Since the method outlined in these patents relies on atmospheric drag experienced by low-earth satellites to greatly simplify the problem, it cannot be used on geosynchronous satellites, which experience no appreciable atmospheric drag. In particular, the presence of atmospheric drag allows the method to know the necessary direction of actuation, the location in the orbit at which to maneuver, and most importantly, adds a large amount of damping to the system. Since geosynchronous spacecraft experience no natural damping, a method for geosynchronous orbital maintenance must correct for orbital errors while also eliminating any oscillations. Since the present invention can function well regardless of the presence or absence of atmospheric drag, the present invention is suitable for orbits of any kind.